The use of high strength fiber-reinforced composite materials in the manufacture of aircraft and other lightweight structures has increased steadily since the introduction of such materials. Composite materials have a high strength-to-weight ratio and stiffness. These properties make composite materials attractive for use in the design of lightweight structures. Some of the drawbacks to using composite materials have been their relatively high fabrication costs, difficulties in manufacturing defect-free parts and poor damage tolerance. Generally, it has been difficult to produce parts formed of high strength composite materials that have the same damage tolerance and fabrication cost as comparable metal parts.
One area of structural and fabrication concern in composite parts is abrupt geometry changes such as the sharp radius of curvation between the webs and caps or flanges commonly found on composite spars, ribs, bulkheads, etc. Generally, in such applications a planar or sine wave shear web is joined to a highly loaded cap or flange at a sharp angle. Difficulties in part fabrication and part geometry often prevent the layers of composite material forming the shear web from extending over the entire width of the caps or flanges. In addition, the caps or flanges generally carry greater loads than the shear webs. Thus, the flanges often include additional reinforcing layers of composite material that are placed over the top of the layers of composite material that form the webs and a portion of the caps or flanges.
Due to the highly loaded nature of the caps or flanges, there is a concern that the loads applied to the caps and flanges will result in separation between the reinforcing cap plies of composite material and the underlying web plies of composite material that are joined to form the flanges. This is of particular concern in composite structures that carry large out of plane loads that tend to pull the caps or flanges away from the shear webs to which they are joined. Composite structures that undergo large out-of-plane loads include aircraft spars, ribs, and bulkheads. Currently, composite structures that undergo large out of plane loads generally use "chicken fasteners" that extend through the cap plies and web plies of composite material to ensure that the layers of composite material remain joined during use of the composite part. For example, composite wing spars and ribs often incorporate chicken fasteners to ensure that the flanges remain joined to the shear webs during loading.
It is desirable in many aircraft applications for the interface between the caps and underlying structure to support pull-off loads on the order of 4,000 to 7,000 lbs/in. Such magnitudes of pull-off loads are greater than are typically achievable using current co-cured composite material technology. Thus, as described above, such highly loaded composite structures incorporate chicken fasteners to increase the maximum pull-off loads.
The use of chicken fasteners in composite parts increases both the weight and fabrication complexity of the finished part. Such fasteners are also an area of concern throughout the maintenance lifetime of the composite part. The fasteners must be frequently inspected to ensure that they do not come loose during the cyclic loading that composite parts undergo.
As can be seen from the discussion above, there exists a need for improved methods of fabricating composite parts that improve pull-off strengths in highly loaded structures such as the flanges or caps of spars, ribs or bulkheads, etc. The present invention is directed towards fulfilling this need.